Aircraft take-off monitoring system



March 1970 R. 1.. HALL. ETAL 3,504,335

I AIRCRAFT TAKE'OFF MONITORING SYSTEM Filed Aug. 14, 1967 3 Sheets-Sheet1 20 LR V 23 22 :I SM X0 F I 24 l IXC I 25 I I svl I XPR T XsvP I I I/zI *2 I v, I V V V I l FIG. 1

2a 27 29 27 ACCEL. ACCEL.

a Q AIRCRAFT I 32 I 3/ \X 35 34 2g w/ osas9o PR m FIG. 2 FIG. 3

I Q 1 2 fg m\ W I I I I d x t 38 gi REAL-TIME I HIGH-SPEED COMPUTER l7KM r-TIME COMPUTER /--""f (SEE FIG. 5) (SEE FIG.6) 7 i I PR 1 -I--'1 Isv| HIGH- SPEED r-TIME COMPUTER svP I y I #s29------+ (SEE F|G.7) 9

4 Fl G //V VENTORS:

ROBERT L. HALL ROLAND H. STEBENS THOMAS COOCH JOSEPH J. ALEKSHUN, JR.

March 31, 1970 L. HALL ET AL 3,504,335

I AIRCRAFT TAKE-OFF MONITORING SYSTEM Filed Aug. 14, 1967 3 Sheets-Sheet2 2 dXp f X -X'r PR 1 +4x V GATE V1 V2 60 a; Y 62 z 2 HOLDER HOLDER I 97I310 92 2 RECYCLE VOLTAGE 2r 11 SOURCE X s XPR FI G. 6

d c T A 77 a! PR 69 7/ 6 dt A I Xsvp E f E 70 74 73 72 75 7a I/VVENTORSIFIG 7 ROBERT L. HALL ROLAND H. SIEBENS THOMAS COOCH JOSEPH J.ALEK$HUN,JR.

ATTORNEYS March 31, 1970 R. L. HALL ET AL 3,504,335

AIRCRAFT TAKB-OFF MONITORING SYSTEM I5 Sheets-Sheet 5 Filed Aug. 14,1967 (LR"XC) X2 d V I E 1/6 720 x s? 1 (LR xc)x2 VARIABLE sfifii'i 1/71/3 I l,- 12/ TAKE-OFF A m VOLTAGE L T SOURCE "9 l jrl 29 (LR"XSM)'(X1svI) VAEILA'ELE (XIJFXSV') 123 AMPLIFIER X PREDICTIVE SW 90 EYE TUBE L xR SM 89 H4 [26 H0 XPR I+ z I v Z XSVP H3 (XPITTXSVP) ED" A R sM)(x +x "RBORT I72 v I30 I V LIGHT I YELLOW ABORT A 6 m 9(lsFs2 LIGHT VOLTAGE E ASOURCE 124 I22 725 FIG. 9

0 I00 CONTROL I+ S'GNAL c I GATE 102 NON CONDUCTIVE L Z GATE I02CONDUCTIVE GATE CONTROL. 0 V-HME VOLTAGE 106 EATE757 707 NON-CONDUCTIVECONTROL SIGNAL 0' A Q GATE I07 E'u'ifi? ER CONDUCTIVE FIG. as

/NVE/V70R$-' ROBERT L. HALL ROLAND H. SIEBENS THOMAS COOCH JOSEPH J.ALEKSHUN,JR.

ATTORNEYS United States Patent Ollice Patented Mar. 31, 1970 Int. Cl.G01c 21/00 US. Cl. 34027 14 Claims ABSTRACT OF THE DISCLOSURE A systemfor monitoring the take-off performance of an aircraft and for providinginformation to the pilot to aid him in his take-off decision in which afirst quantity representing the total distance from the start of takeoffto the aircrafts present position on the runway plus the predicteddistance required to stop from that position from the aircrafts presentvelocity and a second quantity representing the total distance from thestart of take-off to the aircrafts predicted position at which it willhave reached its commitment velocity plus the predicted distance to stopfrom such velocity are each compared with the total safe runway distanceavailable for take-off. In addition, a third quantity representing thedistance from the start of take-off to the aircrafts future position atwhich it will have reached its take-off velocity is compared with thetotal runway distance less the minimum aircraft climb-out distancerequired to clear the end of the runway with a single engine loss.Appropriately displayed information is derived from such comparisons,which are repetitively made with the aid of high speed computers, inorder to indicate as early as possible in the take-off run whether thetake-off should be continued or discontinued.

This invention relates generally to systems for monitoring theperformance of aircraft and, more particularly, to the instrumentationof an improved system for measuring, evaluating and indicating aircraftperformance during the take-off run.

Since runway lengths available for take-off are necessarily limited,members of the flight crew and particularly the pilot of the aircraftmust decide at some point during the take-off run whether to continue orto abort the take-off. Under present take-off procedures the pilot iscalled upon to exercise considerable personal judgment in making suchdecisions based on control panel indications of aircraft operation aswell as on his ability to see the runway and visually estimate theenvironmental conditions and the distances and speeds involved.

Under marginal conditions where the pilot does not have sufficientinformation to make a clear judgment of the situation his decisionbecomes extremely critical. Continuation of the take-off may lead to adisastrous crash while abortion thereof may produce less catastrophicyet serious effects on airport operations and passenger reactions aswell as create possible strain and damage to the aircraft itself. Hence,it is desirable to reduce the necessity for personal judgment of thepilot to a minimum and in cases where such judgment is unavoidable tofurther minimize the possibility of error on his part.

The system of this invention provides the pilot with easily readablepanel indications of the aircrafts present position and certain of theaircrafts predicted future positions under the existing take-offconditions so that a decision to take off or to abort can be made asearly as possible during the take-off run. The system includes anacceleration measuring means and real-time computer means fordetermining the aircrafts present acceleration and for calculating itspresent velocity and position on the runway together with high speedcomputer means for predicting from such real-time information theaircrafts operating velocity and position at future points in time atcritical positions along the runway so that appropriate information formaking a take-off or abort decision can be readily and continuouslydisplayed for pilot observation.

In making such a decision the system takes into account actual andpredicted aircraft operation both prior to and after reaching aspecified commitment velocity conventionally designated as V V isdefined as the velocity at which, if one engine of a multi-engineaircraft should fail, the aircraft either could attain the requiredtake-off height at the end of the runway or could come to a completestop before reaching the end of the take-off surface. The selection ofthe value of such commitment velocity V, depends on the particularaircraft involved and, for any particular aircraft, V can be calculatedprior to take-off in accordance with well-known procedures.

The system of the present invention provides an improved method foraiding the pilot in his take-off decision by deriving informationconcerning present and predicted aircraft positions and runwayconditions primarily based upon an accurate measurement of the aircraftsforward acceleration and supplying him with a clear indication ofwhether to continue or to abort his take-off run based on suchinformation. In such system, a high-speed computer subsystem, utilizinginput quantities obtained from a real-time computer subsystem and fromground calculations made prior to takeoff, predicts the expected valuesof three critical distance quantities. The first quantity represents thetotal distance from the start of take-off to the aircrafts presentposition on the runway plus the predicted distance required to stop fromthat position from the aircrafts present velocity. The second quantityrepresents the total distance from the start of take-off to theaircrafts predicted future position at Which it will have reached thecommitment velocity V plus the predicted distance required to stop fromsuch position from such velocity. The third quantity represents thetotal distance from the start of take-off. to the aircrafts predictedfuture position at which it will have reached its take-off velocitydesignated herein as V The take-off or abort decision is based on acomparison of the first and second distance quantities with the totalsafe runway distance available for take-off, i.e., the total runwaydistance less an arbitrary safety-margin distance, and on a comparisonof the third quantity with the total runway distance less the minimumclimb-out distance required for the aircraft to take off and clearobstacles at the end of the runway under single-engine loss conditions.Such comparisons are discussed in more detail with particular referenceto the Take-off and Abort Conditions (1) through (7) listed below.

Once such comparative relationships are determined, suitableinstrumentation is provided to display the results of such determinationto the pilot in an easily readable fashion. In the system of theinvention some of the parameters used therein are calculated asappropriate constants prior to take-off, others can be continuouslymeasured by available airborne instrumentation and others arecontinuously calculated by airborne real-time computer subsystemequipment or airborne high-speed computer subsystem equipment. Althoughnot necessarily limited thereto, the overall monitoring system can bearranged to be independent of other measurement equipment and systemsalready present in the aircraft and can be made and used as aself-contained, airborne unit not requiring continuous ground-fed datafor its operation.

The specific system and the operation thereof is described morecompletely with the help of the following drawings in which:

FIG. 1 shows a diagrammatic illustration of the various distances andvelocities involved during the aircrafts take-off run along the runway;

FIG. 2 shows a partial cross section of the accelerometer mountingsystem used to measure the aircrafts forward acceleration;

FIG. 3 shows a block diagram of the real-time acceleration measuringsubsystem of the invention;

FIG. 4 shows a simplified block diagram of a real-time computersubsystem and the high-speed computer subsystems of the invention;

FIG. 5 shows a more detailed block diagram of the real-time computersubsystem shown in FIG. 4;

FIG. 6 shows a more detailed block diagram of one of the high-speedcomputer subsystems of FIG. 4;

FIG. 7 shows a more detailed block diagram of another of the high speedcomputer subsystems of FIG. 4;

FIG. 8 shows a partial schematic and partial block diagram of a typicalvalue holder circuit as used in the high-speed computer subsystems ofFIGS. 6 and 7; and

FIG. 9 shows a block diagram of a portion of the overall system whichutilizes output signals from the above computer subsystems to provideappropriate control signals for operating visual panel indicationdevices.

The following distance and velocity definitions will aid inunderstanding the configuration and operation of the system of theinvention. Such quantities and their relationship to the runwaydistances involved are diagrammatically illustrated in FIG. 1 wherein:

L designates the length of the complete runway from a first end linenear the start of the take-off run to a second end line 21 at theopposite end of the runway near the end of the take-off run;

X designates the distance from end line 20 to the point 22 at which theaircraft starts its take-off run (i.e., at time equal to zero);

X designates the distance from end line 21 of the runway to a nearbypoint 23 and represents an arbitrary safety-margin distance;

X designates the minimum distance required for climbout under conditionswhere a multi-engine aircraft is operating under the loss of a singleengine to assure that the aircraft will clear the height of the tallestobstacles located at end line 21 of the runway;

V designates the commitment velocity as defined above;

X designates the distance from end line 20 of the runway to the point 24on the runway at which the velocity V ns reached;

X designates the predicted distance required for the aircraft to come toa stop from a velocity V X designates the distance from end line 20 ofthe runwav to the present position 25 of the aircraft during take-off;

dX /dt designates the present velocity of the aircraft during take-off;

X designates the predicted distance required for the aircraft to come toa stop from the velocity dX /dt;

V designates the take-off velocity of the aircraft; and

X is the distance from end line 20 of the runway to the point 26 on therunway at which the take-off velocity V is reached.

In the system of the invention the basic quantity required to determinethe necessary parameters for indicating aircraft performance is thehorizontal acceleration of the aircraft in the forward direction alongthe runway as the aircraft proceeds from its take-off starting positionat X toward end line 21 thereof. If a single-axis accelerometer is usedto measure such forward acceleration, the accelerometer output will beaffected by pitching, rolling and yawing motions of the aircraft as itproceeds down the runway and by the presence of components ofacceleration due to gravity unless such accelerometer is mounted on asuitable stabilized platform. Such factors tend to introduce measurementerrors into the system which may cause the output signal to lie outsidethe accuracy limits desired for the system. In order to avoid sucherrors in the measurement of acceleration, while at the same timepermitting at least a limited rotation of the aircraft about its pitchand roll axes, and in order to avoid the use of an elaborate andexpensive stabilized platform system, a dual accelerometer system of theconfiguration shown in FIG. 2 can be utilized. The effects of yawingmotions can be eliminated by a proper location of the accelerometers onthe aircraft.

In such structure two accelerometers 27 and 28 are suitably attached toa mounting structure 29 so that their input axes lie in a vertical planealong the center line of the aircraft and are substantiallyperpendicular to each other. A pendulous mass 30 is attached at thebottom of mounting structure 29 and operates to retain the accelerometerinput axes in the vertical plane in the face of rotation of the aircraftabout its roll axis. Since a change in acceleration due to aircraft rollis a function of the cosine of the roll angle, small variations in theroll angle (i.e., the aircraft rolls no more than a few degrees ineither direction from the vertical) produces minimal variations inaccelerometer outputs. Thus, under conditions where the aircraft isexpected to roll only very slightly during the take-off run, pendulousmass 30 may be omitted altogether without producing variations in theaccelerometer output signals outside the desired accuracyspecifications.

Accelerometers 27 and 28 each measure accelerations along the directionsof their input axes, which axes are arranged at installation to form anyappropriate fixed angles with the center line of the aircraft (so longas they are substantially perpendicular to each other, as mentionedabove). During aircraft operation such input axes form angles a and 18,respectively, with the horizontal axis and, consequently, while on and Bvary during takeoff, the quantity (a-l-B) remains essentially equal toIn the drawing the horizontal axis and the aircraft center line areshown as coinciding, although such coincidence is not necessarilymaintained during take-off.

With reference to FIG. 3 the output signal A from accelerometer 27 isfed to the input terminals of a suitable multiplier circuit 31 whichproduces an output signal A representing the square of the accelerometerinput signal. The output signal A from accelerometer 28 is similarly fedto the input terminals of a second multiplier circuit 32 for producingan output signal A These signals are then fed to the input terminals ofa suitable summation amplifier 33 together with a signal g representingthe squareof the acceleration g due to gravity. The output of summationamplifier 33 is then fed as an input signal to a suitable means 34 forobtaining the square root of such input signal, the output of means 34thereby representing the present aircraft horizontal acceleration asmeasured in real-time, so long as the overall average slope of therunway is zero (i.e., the runway is essentially level). When the overallaverage slope of the runway is positive or negative, an appropriatecorrection factor equal to (K g) is introduced as shown. The signal g ismultiplied by K, at coefficient amplifier 35 and then fed to one inputof summation amplifier 36 the other input of which is obtained fromsquare-root means 34. The output of summation amplifier 36 therebyrepresents the acceleration corrected for runway slope characteristics.

The operation of the overall accelerometer computersubsystcm Shown inthe block diagram of FIG. 3 is most easily understood with the help ofthe following equations.

The output signals from accelerometers 27 and 28 may be written asfollows:

[1 sin a+ i (cos a) d X A -g S11] 6- dtz (cos B) where each of thesymbols has been previously defined. Since (ct-H3) equals 90, Equation 2can be rewritten as follows:

A =g cos oz- (sin a) (sin a+COS a) Since (sin Ot+COS a)=l, Equation 4can be simplified as follows:

The values of angles a and B will vary as the aircraft rotates about itspitch axis. However, since the quantities a and ,8 are eliminated inEquation 5, any changes in such angles due to minor pitching of theaircraft will not affect the accuracy of the measurement of aircraftacceleration which then can be expressed in accordance with thefollowing equation:

2 ig V 27 2s (6) The functional mechanization of Equation 6 is shown inthe block diagram of FIG. 3 with an appropriate correction beingsubsequently made at summation amplifier 36 for runway slope asdiscussed above. Further minimization of the errors involved can beachieved by mounting the accelerometers on substantially non-vibratorystructures so that vibrations are reduced.

Thus, FIG. 3 describes a real-time computer subsystem for measuring thebasic quantity required for the overall take-off monitoring system ofthe invention, i.e., a measurement of the horizontal, or forward,acceleration of the aircraft as it proceeds down the runway during itstake-off run.

The overall operation of the remaining computer subsystems of theinvention can be best explained with the help of the block diagramsshown in FIGS. 4, 5, 6, 7 and 8. In describing the functions performedby the mechanization of the system in such figures the followingadditional quantities can be defined:

M designates the mass of the aircraft during take-off;

,u designates the present coeflicient .of rolling friction of theaircraft as calculated in real-time;

W designates the weight of the aircraft and is equivalent to thequantity (Mg);

L designates the lift of the aircraft;

C designates the drag coefiicient of the aircraft;

C designates the lift coefficient of the aircraft;

S designates the wing area of the aircraft;

5 designates the air density;

j U) designates the present engine thrust as a function of time;

,u designates the coefficient of dynamic friction under brakingconditions, sometimes hereinafter referred to as the brakingcoefficient;

1. designates an arbitrary maximum value for such braking coefficient asdiscussed below with reference to Equation 11;

V designates the relative wind velocity in the direction of take-off asmeasured with the aircraft at zero velocity;

dX /d'l generally represents a high-speed computed value of velocity,the use of which is discussed more particularly below;

X generally represents a high-speed computed value of distance, the useof which is discussed more particularly below; and

X designates the high-speed computed value of the distance required toreach the commitment velocity V and X designates the high-speed computedvalue of the distance required to reach the take-off velocity V assumingloss of one engine at V FIG. 4 shows a simplified block diagram of aportion of the overall computer system which is used to calculate therequired distance and velocity qauntities both in realtime by real-timecomputer subsystem 37 and in highspeed time (hereinafter sometimesreferred to as -r-time) by high-speed computer subsystems 38a and 39.Real-time computer 37 receives acceleration input information from thereal-time computer subsystem shown in FIG. 3 as well as informationconcerning present engine thrust f (t), relative wind velocity V andinitial velocity and position quantities V and X respectively. From suchinput signals real-time computer 37 produces output signals representingthe aircrafts present position X and present velocity dX /dt and asignal representing t g. Certain of such signals are appropriately fedto r-time computers 38 and 39 together with the engine thrust andrelative wind velocity signals. Input signals representingground-calculated velocities V and V a ground-calculated arbitrarydistance X for use in recycling the operation of r-time computer 38 asdiscussed in more detail below, and a signal ,u g, rep-resenting anarbitrarily selected maximum value for the coefficient of brakingfriction are also supplied to 'r-ti-me computers 38 and 39 as shown.

The T-time computer 38 is used to compute predicted values in -r-time ofthe distance X from the start of takeoff to the aircrafts predictedposition when it will have reached the commitment velocity V and thedistance X from the start of the take-off to the aircrafts predictedposition when it will have reached its take-off velocity V The r-timecomputer 39 is used to compute the predicted values in T-time of thedistances X and X required for the aircraft to stop from its presentvelocity and from its commitment velocity, respectively. Such quantitiesare dependent on the value of the coefficient of braking friction Whenthe value of a exceeds the arbitrarily selected maximum value r-timecomputer 39 is thereupon arranged to produce such quantities as afunction of the fixed maximum value 1.1. of such coefficient asexplained more fully below.

The above predicted quantities from r-time computers 38 and 39 togetherwith the present distance quantity X produced by real-time computer 37are in turn appropriately summed at summation amplifiers 89 and 90 toproduce the three principal predicted distance quantities required foruse in determining the aircrafts take-off performance. Such quantities,as discussed previously, include (l) the total distance (X +X requiredfor the aircraft to reach its present velocity and to stop therefrom,(2) the predicted total distance (X -+X required for the aircraft toreach its commitment velocity and to stop therefrom, each of suchdistance quantities being ultimately compared with a distance quantity(L X representing the total runway distance less the safety-margindistance as discussed with reference to FIG. 9, and (3) the predicteddistance X required for the aircraft to reach its take off velocitywhich distance quantity is compared with the distance quantity (L Xrepresenting the total runway distance less the minimum distancerequired for climb-out under single-engine loss conditions as alsodiscussed above with reference to Conditions for Continuing Take-OffCondition #1 X23 (LR-X) Condition #2 1 PR Xzs (L -Xe) When V1 Conditionsfor Aborting Take-Off Condition #3 (Normal Abort) Condition #4 (NormalAbort) X XPR+XSVP (L X When Condition #5 (Normal Abort) Condition #6(Normal Abort) Pn+ svPS R sM)} PR X2 (LR XC) When dt Condition #7(Emergency RED Abort) Once suc-h comparisons have been made, suitablereadout or display devices can be actuated to convey such comparativeinformation to the pilot, or other members of the flight crew, asdescribed below with reference to FIG. 9. As discussed more fully withreference to FIG. 9, under Conditions (3) through (6) the pilot stopsthe aircraft by normal braking means while under Condition (7) anemergency or Red abort indicates a possible crash situation under whichthe pilot must take all possible steps (normal braking, reversingengines, etc.) to stop the aircraft immediately.

Before discussing the detailed operation of computers 37, 38 and 39 itis desirable to describe the operational equations which are mechanizedtherein.

The present engine thrust f (t) of an aircraft as a function of time canbe defined by the following equation:

fraf

where the symbols are as defined above and the aircrafts forwardvelocity dX /dt has added thereto a correctional quantity due to therelative wind velocity V In addition, the lift of the aircraft can beexpressed in the following manner:

If Equation 8 is substituted into Equation 7 and the resultingexpression is solved for the acceleration where K is equal to 1 8 M 5D#R'C1.)

written:

d X dX 2 [Hsg+ 2 w+ where K is equal 1 5 M 5 D#s'CL) and, in a mannersimilar to that discussed above, a is a fixed quantity while asdesignates a time variable quantity.

Under practical conditions the coefficients of rolling and brakingfriction, MR and a respectively, are within reasonable limitsessentially proportional to each other so that the following expressioncan be Written:

fls= al a where K is a constant of proportionality. Equation 11 remainsvalid so long as remains within upper and lower limits ps2 and p.respectively.

With reference to FIG. 5, the real-time computer subsystem 37 showntherein is utilized to provide signals representing the aircraftspresent position X its present velocity dX /dt and the quantity mg aspreviously discussed. Such real-tirne quantities are then available forapplication to the T-time, high-speed computer subsystems 38 and 39, theoperations of which are discussed below with reference to FIGS. 6 and 7.

In FIG. 5 the required output quantities are all derived from areal-time measurement of the present forward acceleration d X /a't ofthe aircraft as computed by the accelerometer subsystem shown anddiscussed with reference to FIG. 3. Thus, in FIG. 5 an input signalrepresenting present aircraft forward acceleration is supplied to afirst integrator 41, the output of which is fed to a second integrator42. Input acceleration d X /dt is supplied to integrator 41 through agating circuit 43 which operates so as to produce a real-timecomputation only after the input acceleration exceeds a particularpreselected value. Thus, the acceleration d XPR/dt is simultaneously fedto gate 43 as a gate actuation control signal as shown. Such signal isalso fed to a second gating circuit 44 so that the engine thrust f U)feeds through gate 44 for use in the real-time computation system onlyafter the input acceleration exceeds the above-discussed preselectedvalue. The quantity f O) may be ground calculated as a fixed quantityand fed to computer 37 as a constant value, or it may be derived as apresently measured quantity directly from suitable engineinstrumentation.

Once gates 43 and 44 are actuated to become conduc tive, the inputacceleration is integrated by integrator 41 which produces an outputsignal dX /a't representing the present velocity of the aircraft. Suchvelocity quantity is then fed to integrator 42, the output of whichrepre-- sents the present position X of the aircraft on the runway. Toassure that the present velocity and present position values conform tothe actual initial take-off conditions, the initial velocity V of theaircraft is fed as an initial condition to integrator 41 and the initialposition X of the aircraft on the runway is fed as an initial conditionto integrator 42. If the aircraft is making its take-off from astand-still position, the initial velocity V is equal to zero, while ifthe aircraft is making a running take-off, V will, of course, have aspecific 9 initial value. X and V are illustrated diagrammatically inthe runway diagram of FIG. 1.

The aircraft forward acceleration is also fed simultaneously to acoefiicient amplifier 45 where it is multiplied by the quantity M. Theoutput of coefficient amplifier 45 is fed to one input of a summationamplifier 46, the other input of which is supplied with the enginethrust signal f U) which is fed thereto via gate 44. The output ofsummation amplifier 46 is fed to one input terminal of a secondsummation amplifier 47.

Simultaneously the present aircraft velocity a'X /dt is fed to a firstinput terminal of a summation amplifier 48, the other input of which issupplied with a signal representing the relative wind velocity V whichis fed thereto via gate 43 so that the velocity measurement as used inEquation 8 is thereby suitably corrected for relative wind velocity. Theoutput of summation amplifier 48 is then fed to the inputs of amultiplier circuit 49 so that the output signal from such multiplier isthe square of the input signal thereto. Thus, the quantity is obtainedand supplied to the input of a coefiicient amplifier 50 which multipliessuch expression by the quantity K M. The output of coefiicient amplifier50 is then fed to a second input terminal of summation amplifier 47, theoutput of which is then multiplied by the quantity (-1/M) at coefficientamplifier 51. The output of coefiicient amplifier 51 then represents thequantity (,u g) in accordance with the relationship designated byEquation 9.

Thus, real-time computer 37 as shown in detail in FIG. 5 provides thequantities ,u g, dX /dt and X all of which represent the varying valuesof such quantities as measured in real-time. Such quantities, togetherwith the signal representing (l/Mf (t), as obtained at the output ofcoefiicient amplifier 40, are then fed to the T-time, high speedcomputers 38 and 39 as discussed with reference to FIG. 4 and as shownin more detail in FIGS. 6 and 7.

In FIG. 6 the input quantities (l/M)f (t) and (,u g) are fed to T-timehigh-speed computer 38 via a gating circuit 54. That portion of computer38 comprising integrators 52 and 53, summation amplifiers 55 and 56,multiplier 57 and coefficient amplifier 58 represents a mechanization inT-time of an equation of the form of Equation 9 and produces outputsignals dX /d'f and X which can be appropriately used in the remainingportion of computer 38 to provide output signals X and X; whichrepresent the predicted distances, computed at high-speed in r-time,from the starting end line 20 of the runway to the position the aircraftwill have when it reaches the commitment velocity V and to the positionthe aircraft will have when it reaches its take-off velocity Vrespectively. Such predicted quantities are derived from the real-timequantities representing engine thrust, present velocity, groundcalculated velocities V and V and the variable quantity g.

The input signals (l/M)f (t), a g and a feedback signal related to thevelocity quantity dx /dt are fed to a summation amplifier 55 and, thenceto integrator 52 which produces the integrated signal dX /dr whichrepresents a T-time computed velocity, the significance of which isdiscussed below. That signal is in turn integrated by integrator 53 toproduce an output signal X,, which represents a r-time computerdistance, the significance of which is also discussed below. Thecomputed velocity dx /d-r is added to the relative wind velocity V insummation circuit 56, the output of which is fed to the input terminalsof a multiplier circuit 57 to produce a signal equal to dX, 2 T?) 10This quantity is multiplied by K at coeflicient amplifier 58 and is inturn applied to summation amplifier via gate 54 as a feed back signal.

Gating circuit 54 becomes conductive when its input control signal fromsummation amplifier 59 is positive so that 'r-time computer 38 isoperative only when the output signal X, from integrator 53 is less thanan arbitrarily selected maximum value X which quantities arecontinuously compared at summation amplifier 59 to provide the gatingcontrol signal for gate 54. When the integrated signal X, becomes equalto and subsequently greater than X the output from summation amplifier59 passes through zero and becomes negative and gating circut 54 becomesnon-conductive at which point the integrators recycle to begin theT-time computation again. By suitable control of gate 54 the operationof -r-time computer 38 can be recycled to provide predicted velocity anddistance quantities periodically. The recycling time can be controlledbythe selection of the value of X so that the greater the value of X thelonger the cycling time. For practical operation, X and computerparameters may be appropriately chosen, for example, so that the T-timecomputer 38 is recycled approximately once every second in oneparticular embodiment.

The -r-time computer 38 of FIG. 6, therefore, produces two signals dX/dT and X, at the outputs of integrators 52 and 53, respectively, whichsignals represent predicted -r-time computed velocity and distancequantities.

The computed velocity signal dX.,/d'r is fed to one input of each ofthree summation amplifiers 60, 61 and 62. The aircrafts present velocitydX /dt, as computed in real-time, is fed to a second input of summationamplifier 60. The ground calculated commitment velocity V is fed to asecond input of summation amplifier 61, while the ground calculatedtake-off velocity V is fed to a second input of summation amplifier 62.The outputs of summation amplifiers 60, 61 and 62 provide signals forcontrolling the operation of gating circuits 63, 64 and 65,respectively.

The T-time computed position signal X is fed to the inputs of each ofsaid gates 63, 64 and 65. So long as the T-time computed velocity dX d1-is less than to the realtime computed present aircraft velocity dX /dt,gate 63 is actuated so as to become conductive but value-holder circuit66 produces no output signal, as described more clearly below withreference to FIGS. 8A and 8B. When dX /dT becomes greater than thereal-time computed present velocity, a signal X 'r is produced at theoutput of value holder circuit 66 which value is maintained for use insummation amplifier 88 as shown.

When the T-time computed velocity dX /dT becomes greater than the groundcalculated commitment velocity V value holder circuit 67 similarlyproduces a signal representing the T-time predicted computed distance XThe value of such output signal is maintained at the output ofvalue-holder circuit 67 for use in summation amplifiers 88 and 91 asshown.

When the T-time computed velocity dX /dT becomes greater than the groundcalculated take-off velocity V value-holder circuit 68 similarlyproduces a signal representing the T-time predicted computed distance XThe value of such output signal is maintained at the output ofvalue-holder circuit 68 for use in summation amplifier 91 as shown.

The operations of gates 63, 64, 65, and value-holder circuits 66, 67 and68 are controlled by the output signals from summation amplifiers 60, 61and 62, respectively, as representatively described in more detail laterwith reference to FIGS. 8A and 8B.

:l"he quantities X X and X are supplied to the input terminals ofsummation amplifier 88. The output of summation amplifier 88 thenrepresents the predicted, r-time computed quantity X which is then addedto the predicted, 'r-time computed quantity X at summation amplifier 90(as shown in FIGS. 4 and 9) to produce the quantity (X +X whichrepresents the total predicted -r-time computed distance required forthe aircraft to reach the velocity V plus the distance required to stopfrom such velocity. This output signal is one of the three principaldistance quantities needed to provide appropriate pilot information foruse in accordance with Conditions (1) through (7) as discussed above.

Further, the real-time computed quantity X is added to the predicted,T-time computed quantity X at summation amplifier 89 (as shown in FIGS.4 and 9) to produce the quantity (X +X which represents the totalpredicted, -r-time computed distance required for the aircraft to reachits present velocity plus the distance required to stop from suchvelocity. This output signal is another of the three principalquantities needed to provide appropriate pilot information for use inaccordance with Conditions (1) through (7).

Further, the predicted, r-time computed quantity X is subtracted fromthe predicted, 'r-time computed quantity X at summation amplifier 91.The output of summation amplifier 91 is fed to a coefiicient amplifier92 which multiplies such output by n/n-l, where n is the total number ofengines in the aircraft. For example, for a four-engine aircraft, themultiplier of coefiicient amplifier 92 is equal to 4/3, which operationconverts the output signal from four-engine operation to three-engineoperation so that an adjustment for the difference in acceleration dueto a single engine loss is provided. The output of coefiicient amplifier92 is then compared to the predicted, -time computed quantity X atsummation amplifier 93 to produce an output quantity representing thepredicted, T-time computed distance X required to reach the take-offvelocity V Such output signal is another of the three principalquantities needed to provide appropriate pilot information for use inaccordance with Conditions (1) through (7).

The r-time, high-speed computer subsystem 39 of FIG. 4 is shown in moredetail in FIG. 7 and implements in r-time the braking conditions ofEquation 10. The purpose of computer subsystem 39 is to produce inT-time those quantities which represent the predicted distances, X and Xrequired for the aircraft to stop from its present velocity dX /dt andfrom its commitment velocity V respectively. Such quantities are thenused as input signals to summation amplifiers 89 and 90, the operationof which is described in FIGS. 4 and 9.

In FIG. 7 the stopping distances X and X are each obtained by similarcomputational processes and, as mentioned with reference to FIG. 4,computer 39 requires input signals g and dX dz obtained as real-timecomputed quantities from the operation of real-time computer 37 andinput signal ,u g, V V and K obtained as fixed ground calculatedquantities. The quantity am? is multiplied by a suitable proportionalityconstant K in accordance with Equation 11 to provide the quantity g,where a is the coefiicient of braking friction.

With respect to the computation of X the quantity ,u g is applied to agating circuit 70 via another gating circuit 84, the purpose andoperation of which is described more fully below. The quantity relatedto the computed velocity dX d1- which has been suitably corrected forwind velocity V is also fed to gating circuit 70 after which it iscombined with ,u g in summation amplifier 69. The output of summationamplifier 69 represents a computed acceleration quantity under brakingconditions (more properly described as a deceleration) which isintegrated in integrator 71 to obtain a -r-time computed velocityquantity. This velocity is thereupon subtracted from the real-timecomputed present velocity dX /dz at summation amplifier 75, dX /dt beingfed to amplifier 75 via gating circuit 70. The output of summationamplifier 75 thereupon represents a computed velocity under brakingconditions identified for convenience as dX /dT. This velocity is addedto the relative wind velocity V in summation amplifier 72, the output ofwhich is fed to both input terminals of a multiplier circuit 73 toprovide a quantity Ia-r The latter quantity is multiplied by K atcoefiicient amplifier 74 to produce the appropriate feedback signalmentioned above for summation amplifier 69 via gating circuit 70. Therecycling operation of gate is controlled in a suitable manner discussedin more detail below.

The output dX d1 of summation amplifier 75 which represents a computedvelocity quantity related to the present velocity of the aircraft underbraking conditions is then fed to the input of an integrator circuit 78to produce a 'r-time computed distance quantity under braking conditionswhich is then fed to a value-holder circuit 81 via a gating circuit 77.When the T-time computed velocity signal from integrator 71 becomesgreater than the aircraft real-time computed present velocity (i.e., theoutput of summation amplifier 75 becomes negative), value-holder circuit81 produces an output signal which represents the predicted T-timecomputed value of the distance X required for the aircraft to stop fromits present velocity. The value of such output signal is maintained atsuch value at the output of value holder 81 for use in summationamplifier 89 as shown in FIGS. 4 and 9.

In a similar manner, in order to compute the quantity X the quantity ,ug is fed to a gating circuit 131 whereupon it is combined in summationamplifier 132 With the quantity related to the computed velocity dX /d'rwhich has been suitably corrected for wind velocity V which is also fedto summation amplifier 132 via gating circuit 131. The output ofsummation amplifier 132 represents a computed deceleration quantitywhich is then integrated in integrator 133, the output of whichrepresents a 'r-time computed velocity quantity under brakingconditions. This velocity quantity is subtracted from the groundcalculated commitment velocity V at summation 134, V also being fed tothe input of summation amplifier 134 via gating circuit 131. The outputof summation amplifier 134 thereupon represents a computed velocityquantity under braking conditions identified for convenience as dX /d'n'This velocity is added to the relative wind velocity V in summationamplifier 135, the output of which is fed to both input terminals of amultiplier circuit 136 to provide the quantity d1 2 w'ii The latterquantity is appropriately multiplied by K; at coeflicient amplifier 137to produce the appropriate feedback signal mentioned above for summationamplifier 132 via gating circuit 131.

The output dX /dr of summation amplifier 134 is also fed to anintegrator to produce a -r-time computed distance quantity under brakingconditions which is then fed to a value holder circuit 82 via a gatingcircuit 79. When the r-time computer velocity signal from integrator 133becomes greater than the commitment velocity V (i.e., the output ofsummation amplifier 134 becomes negative), value holder circuit 82produces an output signal which represents the predicted -r-tirnecomputed value of the distance X required for the aircraft to stop fromits commitment velocity. The value of such output signal is maintainedat such value at the output of the value holder circuit 82 for use insummation amplifier as shown in FIGS. 4 and 9.

A signal representing (,u g) is derived from the realtime computation of(u g) by the multiplication thereof by proportionality constant K atcoefiicient amplifier 85. As discussed above, it is necessary to placean upper limit on the value of u For this purpose (u g) is compared withan arbitrarily selected upper limit (,u g) at summation cicruit 87. Theoutput of summation circuit 87 is then used to control the actuation ofgates 83 and 84, the latter being controlled by the negative of thevalue of the signal from circuit 87 as obtained via polarity reversalamplifier 86. As long as (,u g) is less than (,u g), the quantities Xand X are computed by utilizing (u g) as transmitted to gates 70 and 131via gate 84. Whenever (u g) reaches or exceeds its upper limit (u zg),(i.e., the output of summation amplifier 87 goes from a negative to apositive value), gate 83 becomes conductive and gate 84 becomesnon-conductive so thatthe constant quantity (u g) is transmitted togates 70 and 131 via gate 83 for such computation purposes.

The operation of T-time computer 39 is recycled at appropriate times. Inthe computer subsystem as shown in FIG. 7 this recycling operation iscontrolled by the value of the computed velocity quantities aX /d-r anddX /d1- as obtained from the outputs of summation amplifiers 75 and 134,respectively. The output of summation amplifier 75 controls theoperation of gating circuit 138, the output of which in turn controlsthe operation of gating circuit 70. Gate 138 is non-conducting so longas the r-time computed velocity quantity from integrator 71 is less thanthe real-time computed present velocity of the aircraft dX /dt.

Similarly, the output of summation amplifier 134 controls the operationof a gating circuit 139, the output of which in turn controls theoperation of gate 131. Gate 139 is non-conductive so long as thecomputed velocity quantity from integrator 133 is less than thecommitment velocity V The input signals to gates 138 and 139 areobtained from an appropriate recycle voltage source 140.

Thus, T-time computer 39 is recycled (i.e., gates 70 and 131 are madeconductive) only when the computed velocity quantities from integrators71 and 133 become greater than the real-time computed aircraft velocitya'X /dt and the ground calculated commitment velocity V respectively, atwhich times gates 138 and 139 both become conductive and the recyclevoltage is applied to gates 70 and 131. Such operation assures that thecomputer is not recycled until computation of the appropriate distancequantities X and X is completed.

Both T-time computers 38 and 39 utilize unique value holder circuits,such as circuits 66, 67 and 68 in T-time computer 38 and circuits 81 and82 in T-time computer 39. The operation of such circuits can beexplained with the help of the diagrammatic and graphicalrepresentations shown in FIGS. 8A and 8B, respectively which depict theoperation of a typical value holder circuit 100 of the type used in suchsubsystems. The circuit of FIG. 8A shows an input signal A obtained fromthe output of a gate 101, an output signal B, and a control signal Cused for controlling the operation of gate 101 and a. gate 102 which isa part of value holder circuit 100. In FIG. 7, for example, value holdercircuit 82 has a corresponding input signal obtained from gate 79, anoutput signal X and a control signal control signal causes gate 102 tobecome conductive, the voltage value of input signal A is applied viagate 102 to a condenser 104 which thereupon charges up to the valueofsignal A. When gate 101 is subsequently made nonconductive, signal A isno longer applied to condenser 104 but output signal B is retained atthe value of signal A to which such condenser has been charged for aperiod of time dependent on the time constant of the condenser circuit(i.e., the time at which such signal decays to a specified percentage ofits initially charged value). Such time constant may be arranged to berelatively long in comparison to the recycling time of the computersubsystem so that the output of the value holder is maintained at thecondenser-charged value which it achieved after actuation of its controlgate.

FIG. 8B shows a graphical represenation of the relative operations ofgates 101 and 102 where it is assumed that such gates will be in aconductive state when the control signal to the gate is essentiallypositive and in a nonconductive state when the control signal isessentially negative. Thus, as shown in FIG. 8B, when control signal C,designated by line 105, is positive, gate 101 is conductive. As controlsignal C passes through zero and becomes negative, gate 101 remainsconductive only until such signal becomes sufliciently negative, as atthe value designated by point 106, to cause gate 101 to becomenonconductive. Thereafter, so long as control signal C becomes morenegative, gate 101 remains in a nonconductive state.

Line 107 designates the control signal applied to gate 102 andrepresents the negative of control signal C, which has been suitablyreversed in polarity and amplified by amplifier 103. For purposes ofdiscussion, such control signal has been designated as control signal Cin FIGS. 8A and 8B. So long as control signal C is negative, gate 102 isnonconductive. Gate 102' only becomes conductive when control signal Cpasses through zero and reaches a sufficiently positive value, such asthe value shown at point 108, to cause conduction. Gate 102 thereafterremains conductive so long as control signal C continues to become morepositive. Thus, there is a period of time, designated by the time span109 in FIG. 8B, during which both gate 101 and gate 102 aresimultaneously conductive. The duration of this time period isdetermined by the slope of line 107 which can be controlled by the gain(-K of amplifier 103.

Thus, when gate 101 is conductive the output signal A is applied tocapacitor 104 as soon as gate 102 also becomes conductive. Capacitor 104thereupon charges up to the value of signal A and is retained at suchvalue even when such voltage is removed at the time gate 101 becomesnonconductive.

Thus, in FIG. 7, for example, the control signal from summationamplifier 76 is applied to the value holder gate (i.e., the gate, notshown, of value holder circuit 82 equivalent to gate 102 of FIG. 8A)only after a suitable polarity reversal by an appropriate amplifier(also not shown but equivalent to amplifier 103 of FIG. 8A). Thus, thevalue holder gate is controlled by a signal which is the negative(suitably amplified) of the control signal for gate 79. So long as thecomputed velocity dXsl/d'l' is less than V gate 79 is conductive whilethe value-holder gate is nonconductive. At some time after 15 tive isprimarily determined by the gain of the polarity reversal amplifier ofvalue holder circuit 82. Other value holder circuits of FIGS. 6 and 7operate in a similar manner.

Thus, by the use of the acceleration measurement subsystem of FIG. 3,the real-time computer subsystem 37 of FIG. 5, the r-time high speedcomputer subsystems 38 and 39 of FIGS. 6 and 7, respectively, and thesummation amplifiers 89 and 90, the distance quantities shown on theleft-hand sides of the above listed take-off Conditions (1) through (7)are obtained for comparison with the known,'ground-calculated quantitiesshown on the righthand sides of such conditions in order to provide thenecessary information for determining whether take-off should becontinued or aborted. FIG. 9 illustrates how such information can beutilized for visual display to the pilot or to other members of theflight crew.

As shown in FIG. 9, the pilot is given two principal visual indicationsof the performance conditions of the aircraft via the panel devices 110-and 111. Panel device 110 represents a well-known electron-ray tube, orsocalled electric eye or tuning eye indicator tube, designated aspredictive eye tube 110 in FIG. 9. In such a tube a fluorescent pattern113 is caused to open or close in response to an input signal so that afan-like pattern of varying size occurs on the face of the tube. Asdescribed in more detail below, predictive eye tube 110 perates to causethe fluorescent fan pattern 113 to remain open so long as conditionspermitting take-off are maintained. The size of the fluorescent patternopening provides at a glance an indication of the safety margin on whichthe aircraft is operating. For example, if the pattern is open to arelatively large extent and occupies a good portion of the area of theface of the tube, the pilot knows that he is operating well within thepredicted takeoff conditions and has a substantial margin of safety fortake off. As the pattern (the eye) begins to close and, thus, to occupyless and less of the area of the face of the tube, the pilot knows thathe is approaching marginal predicted conditions which indicate that anecessity to abort may occur.

As the fluorescent fan pattern 113 is colored green, a go condition isindicated for a predicted successful take-off. As soon as the eye closes(i.e., the area of the fluorescent pattern 113 is reduced to zero), theface of the tube is caused to glow with an appropriate red color so thatthe tube then indicates to the pilot that a normal abort condition [inaccordance with Conditions (3) through (6) discussed above] exists andhe must take the normal precautions to abort the take-01f. Control ofthe operation of predictive eye tube 110 is discussed more fully below.

The second panel indication provided for the pilot is a RED abort panellight 111. Light 111 is normally in a darkened, or off, condition solong as one of Conditions (1) through (6) exists and either a take-oflsituation or a normal abort situation is indicated. Panel light \111 isonly used when an emergency situation arises which requires the pilot totake all possible steps to stop the aircraft immediately. Suchcircumstances exist under Condition (7) where the distance available fortaking off or for stopping is indicated as being too short to achieveeither a safe take-off or a safe stop, using the normal precautions,once the commitment velocity has been reached. The pilot is thus warnedof possible imminent crash conditions by the actuation of RED abortlight 111 as discussed more fully below.

As shown in FIG. 9, the signals (X +X and (X i-X are obtained at theoutputs of summation amplifiers 89 and 90, respectively, and the signalX is obtained from T-time computer 38. Such signals are applied as shownto one input of summation amplifiers 114, 115 and 116, respectively. Theground calculated quantity (LRXSM) iS fed to the second input terminalsof summation amplifiers 114 and 115, while the ground 16 calculatedquantity (L -X is fed to a second input terminal of summation amplifier116.

The signal from summation amplifier controls the operation of a gatingcircuit 117, the signal from summation amplifier 114 controls theoperation of a gating circuit 118 and the signal from summationamplifier 116 controls the operation of a gating circuit 119. Signalsrepresenting the real-time computed aircraft velocity dX /dt and aground calculated commitment velocity V are fed to the inputs of asummation amplifier 120 to produce an output signal dXpR which is usedto control the operation of a gate 121. The negative of the outputsignal from summation amplifier 120 is used to control a gating circuit122 via polarity reversal amplifier 123. Similarly the negative of theoutput signals from summation amplifiers 114 and 116 are used to controlthe operation of gates 124 and 125, respectively, via polarity reversalamplifiers 126 and 127, respectively.

An appropriate voltage from a voltage source, designated as Take-Offvoltage source 128, is applied to the input of variable gain amplifier129 via a first path through gates .117, 118, 119 and 121 and via secondpath through gates 119 and 122. The gain of the variable gain amplifier129 is controlled in accordance with the value of the output signal,designated as the Variable Margin Signal, from summation amplifier 116.An appropriate voltage from a voltage source, designated as Abortvoltage source 130, is applied to RED abort light 111 via a path throughgates 124, 122 and 125.

The display system shown in FIG. 9 operates in accordance withConditions (1) through (7) as described in more detail below.

In accordance with Condition (1), gates 117, 118, 119 and 121 are allactuated to become simultaneously conductive so that an appropriatevoltage from take-oif voltage source 128 is applied to variable gainamplifier 129 to operate predictive eye tube 110. The area offluorescent pattern 113 is determined by the gain of amplifier 129 whichis in turn controlled by the output of summation amplifier 116, asdiscussed above. Thus, when the quantity (L X is greater than X (i.e.,the takeoff distance available exceeds the predicted take-oil distancerequired by the aircraft), fluorescent pattern 113 is open. The extentto which it opens depends upon how much the available take-off distanceexceeds the predicted take-off distance that is required so that if suchdifference is large (e.g., there is a relatively large amount of runwaydistance available for a safe take-ofl), the eye pattern is open to arelatively large extent. As explained above, the relation between thetake-off distance available and the predicted take-01f distance requiredis indicated to the pilot at a glance by the size of the area offluorescent pattern 113. Under Condition (1) it should be noted thatgates 122, 124 and 125 are non-conducting.

In accordance with Condition (2), gates 119 and .122 are actuated tobecome simultaneously conductive while gate 121 is nonconductive. Undersuch condition a takeoff voltage signal from voltage source 128 is fedto the input of variable gain amplifier 129 via gates 119 and 122 toactuate predictive eye tube 110 to indicate to the pilot that thepredicted conditions are such that take-o'fl? may proceed. Operationunder Condition (2) does not depend on the operation of either gates 117or [118 and the values of the control signals thereto have no particularrelevance under such condition.

Under Condition (3) gates 117 and 122 are nonconductive so that nosignal is applied to variable gain amplifier 128 and the area offluorescent pattern 113 of predictive eye tube 110 is reduced to zero(i.e., the eye is closed). At that time tube 110 is arranged to glowwith a red color to indicate immediately to the pilot that he must takethe normal steps to abort the take-01f.

In a similar manner under Conditions (4) and (5), at least gates 118 or119, respectively, are nonconductive, together with gate 122, so that novoltage is applied to variable gain amplifier 129 and the area offluorescent pattern 113 is closed and predictive eye tube glows red toindicate a condition for normal abortion of the take-off.

Under Condition (6), gate 119 is nonconductive so that no voltage isapplied to variable gain amplifier 129 and a normal abort indication isagain given by predictive eye tube 110. Under such condition, since gate124 is also nonconductive no signal is applied to RED abort light 111and that light remains in a darkened, or off, status.

Under Condition (7), gate 119 is nonconductive but gates 122, 124 and125 are simultaneously actuated to become conductive so that a voltagefrom Abort voltage source 130 is applied to RED abort light 111 so thatit is appropriately lighted. Under such condition both abort light 111and predictive eye tube 110 glow red to indicate to the pilot thatemergency action must be taken to stop the aircraft immediately. Thepilot must then do all he can to avoid or to prepare for a crash (i.e.,reverse engines, alert passengers and crew, etc.).

Thus, in accordance with the operation of the display subsystem of FIG.9 the pilot is provided with an immediate indication as to the predictedconditions under which the aircraft is operating so that he can know notonly whether he should continue his take-off or to abort it but also thegeneral margin of safety under which his take-off is proceeding. Such adisplay system represents only one particular embodiment of a displayfor the pilot or other crew members and other display systems may bedevised by those skilled in the art to accept the input information fromthe appropriate computer subsystems to provide the pilot with anindication as to whether he should continue his take-off or discontinueit.

A further refinement in the display system of FIG. 9 may also be usedwith reference to a caution or yellow panel light 112 shown therein. Asmentioned above with reference to FIG. 7, it is often desirable toindicate when the braking coefficient of friction as exceeds anarbitrarily selected maximum value p. Such a condition is indicated bythe yellow panel light 112. which is actuated by the signal g(u ,u z)from the output of summation amplifier 87 in FIG. 7. So long as as isless than p. such signal is negative and yellow caution light 112 is ina darkened or off condition. As soon as u exceeds the input controlsignal to yellow light 112 becomes positive and the light is turned onto indicate such a condition to the pilot. From such indication thepilot is immediately made aware that his runway conditions or theconditions of the aircraft (brake dragging, improper tire inflation,etc.) are different than those assumed prior to start of the take-offrun. If, for example, predictive eye tube 110 shows minimal marginaloperation, the actuation of yellow panel light 112 and the observationof external weather conditions by the pilot will provide him with abetter assessment of or feel for the reasons behind such minimalmarginal performance. If the take-off is ultimately aborted, theexistence and time of such an indication may provide a significant clueas to the cause of the aborted take-off.

Other modifications to the overall system of the invention may occur tothose skilled in the art without departing from the spirit and scope ofthe invention. Hence, the invention is not to be misconstrued as limitedto the particular embodiment as described herein except as defined bythe appended claims.

What is claimed is:

1. An aircraft take-01f monitoring system comprising means for comparingthe predicted distance required for said aircraft to reach its presentposition and to stop from said position from its present velocity withthe safe runway distance available for take-off; means for comparing thepredicted distance required for said aircraft to reach a position whereit will have attained a preselected commitment velocity and to stop fromsaid position from said commitment velocity with said safe runwaydistance available for take-off;

means for comparing the predicted distance required for said aircraft toreach a position Where it will have attained its take-off velocity withthe total runway distance less the minimum distance required for saidaircraft to take off with the loss of single engine and still clear apreselected height at the end of said runway; and

means responsive to each of said comparing means for providing operatinginformation to at least one mem-' ber of the crew of said aircraftduring the take-off run of said aircraft. 2. An aircraft take-offmonitoring system comprising means for measuring theforward accelerationof an aircraft during its take-off run along a runway;

real-time computer means responsive to said acceleration measurement forcomputing the present velocity of said aircraft and the distancetraveled by said aircraft to reach its present position on said runy;

high speed computer means responsive to said computed present velocityand distance for computing predicted values of a first quantity equal tothe distance required for said aircraft to reach its present positionplus the distance required to stop from said position from said presentvelocity, asecond quantityequal to the distance required for saidaircraft to reach a position where it will have attained a preselectedcommitment velocity plus the distance required to stop from saidposition from said commitment velocity, and a third quantity equal tothe distance required for said aircraft to reach a position where itwill have attained its take-off velocity;

means for comparing said first and second quantities with the total saferunway distance available for take-off and for comparing said thirdquantity with a distance equal to the total runway distance less theminimum climb-out distance required for said aircraft to take off withthe loss of a single engine and still clear a preselected height at theend of said runway; and

means responsive to said comparing means for displaying operatinginformation to the crew of said aircraft during the take-off run of saidaircraft.

3. An aircraft take-off monitoring ssytem in accordance with claim 2wherein said computing means includes means for correcting saidacceleration measurement for the effects of pitching motions on saidaircraft during its take-off run.

4. An aircraft take-off monitoring system in accordance with claim 3wherein said computing means further includes means for correcting saidacceleration measurement for the effects of the acceleration of saidaircraft due to gravity during its take-off run.

5. An aircraft take-off monitoring system in accordance with claim 3wherein said gravity correction means further includes means foradjusting said gravity computation in accordance with the average slopeof the runway on which said aircraft is making its take-off run.

6. An aircraft take-off monitoring system in accordance with claim 2wherein said high speed computer means includes a first high speedcomputer for computing pedicted output values of the distance traveledby said aircraft to reach its present position, the distance saidaircraft will have traveled when it reaches its commitment velocity andthe distance said aircraft will have traveled when it reaches itstake-off velocity;

a second high speed computer for computing the predicted output valuesof the distance required for said aircraft to stop at its presentvelocity and the 19 distance required for said aircraft to stop at itscommitment velocity; and

means responsive to said real-time and to said first and second highspeed computer means for determining said first, second and thirdquantities.

7. An aircraft take-off monitoring system in accordance with claim 6wherein said predicted output values computed by said first and saidsecond computers are corrected for the effects of wind velocity.

8. An aircraft take-off monitoring system in accordance with claim 6wherein said first and said second computer means further include aplurality of gating circuits; and

means for controlling the operation of said gating circuits to causesaid computer means to compute said predicted output values cyclicallyat a controllable rate.

9. An aircraft take-off monitoring system in accordance with claim 2wherein said display means includes an indicator means adapted to be paced in a first state to indicate that take-off should be continued andadapted to be placed in a second state to indicate that said takeoffshould be aborted.

10. An aircraft take-01f monitoring system in accordance with claim 9wherein said indicator means includes means for further providing anindication of the margin of operating safety of said aircraft in saidfirst state.

11. An aircraft take-off monitoring system in accordance with claim 10wherein said indicator means comprises an electron-ray tube having afluorescent pattern the area of which is controlled in response to thecomparison of said third quantity with said total runway distance lesssaid climb-out distance.

12. An aircraft take-off monitoring system in accordance with claim 9wherein said display means further includes means for indicatingconditions in which emergency stopping action must be taken to abort thetakeoff of said aircraft.

13. An aircraft take-off monitoring system in accordance with claim 9wherein said display means further includes means for indicating whenthe coefiicient of braking friction of said aircraft exceeds apredetermined maximum value.

14. An aircraft take-off monitoring system comprismg means for obtaininga first quantity representing the distance traveled by said aircraft toreach its present position on a runway plus the predicted distancerequired to stop from said position from the present velocity of saidaircraft;

means for obtaining a second quantity representing the predicteddistance required for said aircraft to reach a position on said runwaywhen it will have attained a preselected commitment velocity plus thepredicted distance required to stop from said position from saidcommitment velocity;

means for obtaining a third quantity representing the predicted distancerequired for said aircraft to reach a position on said runway where itwill have attained its take-off velocity; means for indicating thattake-off should be continued (a) when during the time period when thepresent velocity of said aircraft is less than or equal to saidcommitment velocity each of said first and said second quantities isless than or equal to a fourth quantity representing the total saferunway distance available for take-01f and said third quantity is lessthan or equal to a fifth quantity representing the total runway distanceless the minimum climb-out distance required for said aircraft to takeoff with the loss of a single engine and still clear a pre-selectedheight at the end of said runway; and ('b) when during the time periodwhen said present velocity is greater than said commitment velocity saidthird quantity is less than or equal to said fifth quantity;

means for indicating that said take-ofi should be dis- 2 continued (a)when during the time period when said present velocity is less than orequal to said commitment velocity either one of said first or saidsecond quantities is greater than said fourth quantity;

(b) when during the time period when said present velocity is less thanor equal to said commitment velocity said third quantity is greater thansaid fifth quantity; and

(0) when during the time period when said present velocity is greaterthan said commitment velocity said first quantity is less than or equalto said fourth quantity and said third quantity is greater than saidfifth quantity; and

means for indicating that emergency action should be taken todiscontinue said take-01f when during the time period when said presentvelocity is greater than said commitment velocity said first quantity isgreater than said fourth quantity and said third quantity is greaterthan said fifth quantity.

References Cited UNITED STATES PATENTS 3,048,329 8/1962 Berggren235150.22 3,077,110 2/1963 Gold 235-150.22 XR 3,111,577 11/1963 DeGraffenried 34027 XR ALVIN H. WARINGPrimary Examiner U.S. Cl. X.R.

